A propellant for solid rocket motors normally comprises a propellant grain block housed inside a tubular structure with optional interposition of a heat protection, and front and rear walls connected to the tubular structure. The assembly formed by the tubular structure and the two walls constitutes the combustion chamber.
Various methods are currently used for producing solid propellant rocket motors.
Some of these methods consist first of all in producing the tubular structure of the chamber and then in casting the propellant block therein. This is the case with rocket motors in metal or in composite material. The assembly can be completed with detachable end walls wedged in position as described in French Patent application No. 86 00 877 (Publication No. 2 593 238). The chamber obtained according to this technique is rigid and able to withstand high temperatures due to the fact that its constituting elements are produced separately and that in such a case the requirements for each one can be met.
However, it is very difficult with this particular technique to introduce, in the tubular structure of the combustion chamber, the propellant which is constituted of a block or, in some cases, of an assembly of segments, because this introduction ends up in a groping assembly operation.
French Patent No. 1 386 856 did propose to produce the combustion chamber in modules form, each module comprising a tubular sector forming one part of the tubular structure , and a cylindrical sector forming one part of the propellant block and being fixed on the tubular sector. In the case of rocket motors of large dimensions, the object was to convey the modules separately to the launching site, to be thereafter assembled , which avoided having to transport the fully mounted propelling chamber. Undoubtedly, this method raises problems which are difficult to solve, as regards both the manufacture of the modules and their assembly in order to obtain, by simple joining-up of the modules, a chamber which has a defect-free propellant block as well as a faultless tubular structure.
Another known method, such as that described in French Patent No. 1 356 673 or, more recently, in French Patent No. 83 15 263 (Publication No. 2 552 494), consists in first producing the propellant, and then producing the tubular structure of the combustion chamber in a filamentary composite material, such as of the glass/epoxy resin type or of the carbon/epoxy resin type, by winding the material on the propellant and polymerizing the resin. This technique, however, requires the use of a resin having a relatively low polymerization temperature, i.e. less than about 100.degree. C., because of the presence of the propellant. It becomes then difficult, if not impossible, to obtain a structure which will be perfectly capable of withstanding the heat which the missile containing the rocket motor is subjected to.